The subject matter disclosed herein relates to gas turbine systems, and more particularly to a thermally actuated assembly for controlling a cooling airflow path.
Gas turbine systems include numerous areas that are temperature sensitive based on component materials and efficiency improving schemes. Such areas are often sectioned off and may be provided with a cooling source to ensure proper temperature regulation to maintain service life and improve efficiency of the overall gas turbine system. A compressor of the gas turbine system is often the cooling source and any flow extracted from the compressor to serve a cooling function detracts from the amount of flow that is delivered from the compressor to a turbine for work that is converted into energy. Such detractions are considered chargeable flow losses and reduction of these losses is desirable.
An example of a temperature sensitive area of the gas turbine system is proximate the rotor and turbine blade disks in operable connection thereto. A rim cavity is often included proximate such areas and requires a cooling flow for purging of hot gas from a hot gas path that travels at a relatively radially outward location over stator vanes and turbine blades. A seal, such as a brush seal is typically included proximate the rotor and within a path leading to the rim cavity, however, the seal gradually wears away over the service life of the gas turbine system and a greater volumetric flow rate of cooling air from the cooling source continuously enters the rim cavity during the wearing process. To accommodate the lower volumetric flow rate passing through the path earlier in the service life of the seal, a cooling flow passage is included to allow cooling flow to reach the rim cavity. As the seal wears away, an unnecessarily high amount of cooling flow reaches the rim cavity and overall gas turbine efficiency is decreased.